A fan drive gear system for driving a fan in a gas turbine engine having a high bypass ratio

ABSTRACT

A gas turbine engine includes a fan section that includes a fan rotatable about an axis of rotation of the gas turbine engine. A speed reduction device is connected to the fan. The speed reduction device includes a star drive gear system with a star gear ratio of at least 1.5. A bypass ratio is greater than about 11.0.

CROSS-REFERENCE TO RELATED APPLICATIONS

This disclosure is a continuation in part of PCT/US2013/061115 filed onSep. 23, 2013, which claims priority to U.S. Provisional PatentApplication No. 61/706,212 filed on Sep. 27, 2012.

BACKGROUND

This disclosure relates to a gas turbine engine, and more particularlyto a method for setting a gear ratio of a fan drive gear system of a gasturbine engine.

A gas turbine engine may include a fan section, a compressor section, acombustor section, and a turbine section. Air entering the compressorsection is compressed and delivered into the combustor section where itis mixed with fuel and ignited to generate a high-speed exhaust gasflow. The high-speed exhaust gas flow expands through the turbinesection to drive the compressor and the fan section. Among othervariations, the compressor section can include low and high pressurecompressors, and the turbine section can include low and high pressureturbines.

Typically, a high pressure turbine drives a high pressure compressorthrough an outer shaft to form a high spool, and a low pressure turbinedrives a low pressure compressor through an inner shaft to form a lowspool. The fan section may also be driven by the inner shaft. A directdrive gas turbine engine may include a fan section driven by the lowspool such that a low pressure compressor, low pressure turbine, and fansection rotate at a common speed in a common direction.

A speed reduction device, which may be a fan drive gear system or othermechanism, may be utilized to drive the fan section such that the fansection may rotate at a speed different than the turbine section. Thisallows for an overall increase in propulsive efficiency of the engine.In such engine architectures, a shaft driven by one of the turbinesections provides an input to the speed reduction device that drives thefan section at a reduced speed such that both the turbine section andthe fan section can rotate at closer to optimal speeds.

Although gas turbine engines utilizing speed change mechanisms aregenerally known to be capable of improved propulsive efficiency relativeto conventional engines, gas turbine engine manufacturers continue toseek further improvements to engine performance including improvementsto thermal, transfer and propulsive efficiencies.

SUMMARY

In one exemplary embodiment, a gas turbine engine includes a fan sectionthat includes a fan rotatable about an axis of rotation of the gasturbine engine. A speed reduction device is connected to the fan. Thespeed reduction device includes a star drive gear system with a stargear ratio of at least 1.5. A bypass ratio is greater than about 11.0.

In a further embodiment of the above, the speed reduction deviceincludes a star gear system gear ratio of at least 2.6.

In a further embodiment of any of the above, the speed reduction deviceincludes a system gear ratio less than or equal to 4.1.

In a further embodiment of any of the above, the bypass ratio is lessthan about 22.0.

In a further embodiment of any of the above, the fan blade tip speed ofthe fan section is greater than about 1000 ft/sec and less than about1400 ft/sec.

In a further embodiment of any of the above, the star system includes asun gear, a plurality of star gears, a ring gear, and a carrier.

In a further embodiment of any of the above, each of the plurality ofstar gears includes at least one bearing.

In a further embodiment of any of the above, the carrier is fixed fromrotation.

In a further embodiment of any of the above, a low pressure turbine ismechanically attached to the sun gear.

In a further embodiment of any of the above, a fan section ismechanically attached to the ring gear.

In a further embodiment of any of the above, an input of the speedreduction device is rotatable in a first direction and an output of thespeed reduction device is rotatable in a second direction opposite tothe first direction.

In a further embodiment of any of the above, a low pressure turbinesection is in communication with the speed reduction device. The lowpressure turbine section includes at least three stages.

In a further embodiment of any of the above, there is a low pressureturbine which is one of three turbine rotors. The low pressure turbinedrives the fan section and the other two of the three turbine rotorseach drive a compressor section.

In a further embodiment of any of the above, there is a high pressureturbine, with each of a low pressure turbine and the high pressureturbine driving a compressor rotor.

In a further embodiment of any of the above, the speed reduction deviceis positioned intermediate a compressor rotor driven by the low pressureturbine and the fan section.

In a further embodiment of any of the above, the speed reduction deviceis positioned intermediate the low pressure turbine and the compressorrotor driven by the low pressure turbine.

In another exemplary embodiment, a fan drive gear module for a gasturbine engine includes a star drive gear system with a speed reductionratio of at least 1.5. The star drive gear system is for driving a gasturbine engine with a bypass ratio greater than about 11.0.

In a further embodiment of any of the above, the speed reduction ratiois greater than about 2.6 and less than or equal to 4.1.

In a further embodiment of any of the above, the star drive gear systemis configured to drive a fan section with a fan blade tip speed greaterthan about 1000 ft/sec and less than about 1400 ft/sec.

In a further embodiment of any of the above, the bypass ratio is lessthan about 22.0.

In a further embodiment of any of the above, there is a low pressureturbine which is one of three turbine rotors. The low pressure turbinedrives a fan section and the other two of the three turbine rotors eachdrive a compressor section.

In a further embodiment of any of the above, there is a high pressureturbine, with each of a low pressure turbine and the high pressureturbine driving a compressor rotor.

In a further embodiment of any of the above, the speed reduction deviceis positioned intermediate a compressor rotor driven by the low pressureturbine and a fan section.

In a further embodiment of any of the above, the speed reduction deviceis positioned intermediate the low pressure turbine and the compressorrotor driven by the low pressure turbine.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates a schematic, cross-sectional view of an example gasturbine engine.

FIG. 2 illustrates a schematic view of one configuration of a low speedspool that can be incorporated into a gas turbine engine.

FIG. 3 illustrates a fan drive gear system that can be incorporated intoa gas turbine engine.

FIG. 4 shows another embodiment.

FIG. 5 shows yet another embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The exemplarygas turbine engine 20 is a two-spool turbofan engine that generallyincorporates a fan section 22, a compressor section 24, a combustorsection 26 and a turbine section 28. Alternative engines might includean augmentor section (not shown) among other systems or features. Thefan section 22 drives air along a bypass flow path B, while thecompressor section 24 drives air along a core flow path C forcompression and communication into the combustor section 26. The hotcombustion gases generated in the combustor section 26 are expandedthrough the turbine section 28. Although depicted as a two-spoolturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto two-spool turbofan engines and these teachings could extend to othertypes of engines, including but not limited to, three-spool enginearchitectures.

The exemplary gas turbine engine 20 generally includes a low speed spool30 and a high speed spool 32 mounted for rotation about an enginecenterline longitudinal axis A. The low speed spool 30 and the highspeed spool 32 may be mounted relative to an engine static structure 33via several bearing systems 31. It should be understood that otherbearing systems 31 may alternatively or additionally be provided, andthe location of bearing systems 31 may be varied as appropriate to theapplication.

The low speed spool 30 generally includes an inner shaft 34 thatinterconnects a fan 36, a low pressure compressor 38 and a low pressureturbine 39. The inner shaft 34 can be connected to the fan 36 through aspeed change mechanism, which in exemplary gas turbine engine 20 isillustrated as a geared architecture 45, such as a fan drive gear system50 (see FIGS. 2 and 3). The speed change mechanism drives the fan 36 ata lower speed than the low speed spool 30. The high speed spool 32includes an outer shaft 35 that interconnects a high pressure compressor37 and a high pressure turbine 40. In this embodiment, the inner shaft34 and the outer shaft 35 are supported at various axial locations bybearing systems 31 positioned within the engine static structure 33.

A combustor 42 is arranged in exemplary gas turbine 20 between the highpressure compressor 37 and the high pressure turbine 40. A mid-turbineframe 44 may be arranged generally between the high pressure turbine 40and the low pressure turbine 39. The mid-turbine frame 44 can supportone or more bearing systems 31 of the turbine section 28. Themid-turbine frame 44 may include one or more airfoils 46 that extendwithin the core flow path C. It will be appreciated that each of thepositions of the fan section 22, compressor section 24, combustorsection 26, turbine section 28, and fan drive gear system 50 may bevaried. For example, gear system 50 may be located aft of combustorsection 26 or even aft of turbine section 28, and fan section 22 may bepositioned forward or aft of the location of gear system 50.

The inner shaft 34 and the outer shaft 35 are concentric and rotate viathe bearing systems 31 about the engine centerline longitudinal axis A,which is co-linear with their longitudinal axes. The core airflow iscompressed by the low pressure compressor 38 and the high pressurecompressor 37, is mixed with fuel and burned in the combustor 42, and isthen expanded over the high pressure turbine 40 and the low pressureturbine 39. The high pressure turbine 40 and the low pressure turbine 39rotationally drive the respective high speed spool 32 and the low speedspool 30 in response to the expansion.

In a non-limiting embodiment, the gas turbine engine 20 is a high-bypassgeared aircraft engine. In a further example, the gas turbine engine 20bypass ratio is greater than about six (6:1). The geared architecture 45can include an epicyclic gear train, such as a planetary gear system, astar gear system, or other gear system. The geared architecture 45enables operation of the low speed spool 30 at higher speeds, which canenable an increase in the operational efficiency of the low pressurecompressor 38 and low pressure turbine 39, and render increased pressurein a fewer number of stages.

The pressure ratio of the low pressure turbine 39 can be pressuremeasured prior to the inlet of the low pressure turbine 39 as related tothe pressure at the outlet of the low pressure turbine 39 and prior toan exhaust nozzle of the gas turbine engine 20. In one non-limitingembodiment, the bypass ratio of the gas turbine engine 20 is greaterthan about ten (10:1), the fan diameter is significantly larger thanthat of the low pressure compressor 38, and the low pressure turbine 39has a pressure ratio that is greater than about five (5:1). In anothernon-limiting embodiment, the bypass ratio is greater than 11 and lessthan 22, or greater than 13 and less than 20. It should be understood,however, that the above parameters are only exemplary of a gearedarchitecture engine or other engine using a speed change mechanism, andthat the present disclosure is applicable to other gas turbine engines,including direct drive turbofans. In one non-limiting embodiment, thelow pressure turbine 39 includes at least one stage and no more thaneight stages, or at least three stages and no more than six stages. Inanother non-limiting embodiment, the low pressure turbine 39 includes atleast three stages and no more than four stages.

In this embodiment of the exemplary gas turbine engine 20, a significantamount of thrust is provided by the bypass flow path B due to the highbypass ratio. The fan section 22 of the gas turbine engine 20 isdesigned for a particular flight condition—typically cruise at about 0.8Mach and about 35,000 feet. This flight condition, with the gas turbineengine 20 at its best fuel consumption, is also known as bucket cruiseThrust Specific Fuel Consumption (TSFC). TSFC is an industry standardparameter of fuel consumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fansection 22 without the use of a Fan Exit Guide Vane system. The low FanPressure Ratio according to one non-limiting embodiment of the examplegas turbine engine 20 is less than 1.45. In another non-limitingembodiment of the example gas turbine engine 20, the Fan Pressure Ratiois less than 1.38 and greater than 1.25. In another non-limitingembodiment, the fan pressure ratio is less than 1.48. In anothernon-limiting embodiment, the fan pressure ratio is less than 1.52. Inanother non-limiting embodiment, the fan pressure ratio is less than1.7. Low Corrected Fan Tip Speed is the actual fan tip speed divided byan industry standard temperature correction of [(Tram ° R)/(518.7°R)]^(0.5), where T represents the ambient temperature in degreesRankine. The Low Corrected Fan Tip Speed according to one non-limitingembodiment of the example gas turbine engine 20 is less than about 1150fps (351 m/s). The Low Corrected Fan Tip Speed according to anothernon-limiting embodiment of the example gas turbine engine 20 is lessthan about 1400 fps (427 m/s). The Low Corrected Fan Tip Speed accordingto another non-limiting embodiment of the example gas turbine engine 20is greater than about 1000 fps (305 m/s).

FIG. 2 schematically illustrates the low speed spool 30 of the gasturbine engine 20. The low speed spool 30 includes the fan 36, the lowpressure compressor 38, and the low pressure turbine 39. The inner shaft34 interconnects the fan 36, the low pressure compressor 38, and the lowpressure turbine 39. The inner shaft 34 is connected to the fan 36through the fan drive gear system 50. In this embodiment, the fan drivegear system 50 provides for counter-rotation of the low pressure turbine39 and the fan 36. For example, the fan 36 rotates in a first directionD1, whereas the low pressure turbine 39 rotates in a second direction D2that is opposite of the first direction D1.

FIG. 3 illustrates one example embodiment of the fan drive gear system50 incorporated into the gas turbine engine 20 to provide forcounter-rotation of the fan 36 and the low pressure turbine 39. In thisembodiment, the fan drive gear system 50 includes a star gear systemwith a sun gear 52, a ring gear 54 disposed about the sun gear 52, and aplurality of star gears 56 having journal bearings 57 positioned betweenthe sun gear 52 and the ring gear 54. A fixed carrier 58 carries and isattached to each of the star gears 56. In this embodiment, the fixedcarrier 58 does not rotate and is connected to a grounded structure 55of the gas turbine engine 20.

The sun gear 52 receives an input from the low pressure turbine 39 (seeFIG. 2) and rotates in the first direction D1 thereby turning theplurality of star gears 56 in a second direction D2 that is opposite ofthe first direction D1. Movement of the plurality of star gears 56 istransmitted to the ring gear 54 which rotates in the second direction D2opposite from the first direction D1 of the sun gear 52. The ring gear54 is connected to the fan 36 for rotating the fan 36 (see FIG. 2) inthe second direction D2.

A star system gear ratio of the fan drive gear system 50 is determinedby measuring a diameter of the ring gear 54 and dividing that diameterby a diameter of the sun gear 52. In one embodiment, the star systemgear ratio of the geared architecture 45 is between 1.5 and 4.1. Inanother embodiment, the system gear ratio of the fan drive gear system50 is between 2.6 and 4.1. When the star system gear ratio is below 1.5,the sun gear 52 is relatively much larger than the star gears 56. Thissize differential reduces the load the star gears 56 are capable ofcarrying because of the reduction in size of the star gear journalbearings 57. When the star system gear ratio is above 4.1, the sun gear52 may be much smaller than the star gears 56. This size differentialincreases the size of the star gear 56 journal bearings 57 but reducesthe load the sun gear 52 is capable of carrying because of its reducedsize and number of teeth. Alternatively, roller bearings could be usedin place of journal bearings 57.

Improving performance of the gas turbine engine 20 begins by determiningfan tip speed boundary conditions for at least one fan blade of the fan36 to define the speed of the tip of the fan blade. The maximum fandiameter is determined based on the projected fuel burn derived frombalancing engine efficiency, mass of air through the bypass flow path B,and engine weight increase due to the size of the fan blades.

Boundary conditions are then determined for the rotor of each stage ofthe low pressure turbine 39 to define the speed of the rotor tip and todefine the size of the rotor and the number of stages in the lowpressure turbine 39 based on the efficiency of low pressure turbine 39and the low pressure compressor 38.

Constraints regarding stress levels in the rotor and the fan blade areutilized to determine if the rotary speed of the fan 36 and the lowpressure turbine 39 will meet a desired number of operating life cycles.If the stress levels in the rotor or the fan blade are too high, thegear ratio of the fan drive gear system 50 can be lowered and the numberof stages of the low pressure turbine 39 or annular area of the lowpressure turbine 39 can be increased.

FIG. 4 shows an embodiment 100, wherein there is a fan drive turbine 108driving a shaft 106 to in turn drive a fan rotor 102. A gear reduction104 may be positioned between the fan drive turbine 108 and the fanrotor 102. This gear reduction 104 may be structured and operate likethe geared architecture 45 disclosed above. A compressor rotor 110 isdriven by an intermediate pressure turbine 112, and a second stagecompressor rotor 114 is driven by a turbine rotor 116. A combustionsection 118 is positioned intermediate the compressor rotor 114 and theturbine section 116.

FIG. 5 shows yet another embodiment 200 wherein a fan rotor 202 and afirst stage compressor 204 rotate at a common speed. The gear reduction206 (which may be structured as the geared architecture 45 disclosedabove) is intermediate the compressor rotor 204 and a shaft 208 which isdriven by a low pressure turbine section.

Although the different non-limiting embodiments are illustrated ashaving specific components, the embodiments of this disclosure are notlimited to those particular combinations. It is possible to use some ofthe components or features from any of the non-limiting embodiments incombination with features or components from any of the othernon-limiting embodiments.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed and illustrated in these exemplary embodiments,other arrangements could also benefit from the teachings of thisdisclosure.

The foregoing description shall be interpreted as illustrative and notin any limiting sense. A worker of ordinary skill in the art wouldunderstand that certain modifications could come within the scope ofthis disclosure. For these reasons, the following claim should bestudied to determine the true scope and content of this disclosure.

What is claimed is:
 1. A gas turbine engine comprising: a fan sectionincluding a fan rotatable about an axis of rotation of the gas turbineengine; and a speed reduction device connected to the fan, wherein thespeed reduction device includes a star drive gear system with a stargear ratio of at least 1.5, wherein a bypass ratio is greater than about11.0.
 2. The gas turbine engine of claim 1, wherein the speed reductiondevice includes a star gear system gear ratio of at least 2.6.
 3. Thegas turbine engine of claim 2, wherein the speed reduction deviceincludes a system gear ratio less than or equal to 4.1.
 4. The gasturbine engine of claim 3, including the bypass ratio is less than about22.0.
 5. The gas turbine engine of claim 4, wherein the fan blade tipspeed of the fan section is greater than about 1000 ft/sec and less thanabout 1400 ft/sec.
 6. The gas turbine engine of claim 1, wherein thestar system includes a sun gear, a plurality of star gears, a ring gear,and a carrier.
 7. The gas turbine engine of claim 6, wherein each of theplurality of star gears includes at least one bearing.
 8. The gasturbine engine of claim 7, wherein the carrier is fixed from rotation.9. The gas turbine engine of claim 8, wherein a low pressure turbine ismechanically attached to the sun gear.
 10. The gas turbine engine ofclaim 9, wherein a fan section is mechanically attached to the ringgear.
 11. The gas turbine engine of claim 1, wherein an input of thespeed reduction device is rotatable in a first direction and an outputof the speed reduction device is rotatable in a second directionopposite to the first direction.
 12. The gas turbine engine of claim 11,including a low pressure turbine section in communication with the speedreduction device, wherein the low pressure turbine section includes atleast three stages.
 13. The gas turbine engine as recited in claim 1,further comprising a low pressure turbine which is one of three turbinerotors, and the low pressure turbine drives the fan section and theother two of the three turbine rotors each drive a compressor section.14. The gas turbine engine as recited in claim 1, further comprising ahigh pressure turbine, with each of a low pressure turbine and the highpressure turbine driving a compressor rotor.
 15. The gas turbine engineas recited in claim 14, wherein the speed reduction device is positionedintermediate a compressor rotor driven by the low pressure turbine andthe fan section.
 16. The gas turbine engine as recited in claim 14,wherein the speed reduction device is positioned intermediate the lowpressure turbine and the compressor rotor driven by the low pressureturbine.
 17. A fan drive gear module for a gas turbine enginecomprising: a star drive gear system with a speed reduction ratio of atleast 1.5, wherein the star drive gear system for driving a gas turbineengine with a bypass ratio greater than about 11.0.
 18. The fan drivegear module of claim 17, wherein the speed reduction ratio is greaterthan about 2.6 and less than or equal to 4.1.
 19. The fan drive gearmodule of claim 17, wherein the star drive gear system is configured todrive a fan section with a fan blade tip speed greater than about 1000ft/sec and less than about 1400 ft/sec.
 20. The fan drive gear module ofclaim 17, wherein the bypass ratio is less than about 22.0.
 21. The fandrive gear module as recited in claim 17, further comprising a lowpressure turbine which is one of three turbine rotors, and the lowpressure turbine drives a fan section and the other two of the threeturbine rotors each drive a compressor section.
 22. The fan drive gearmodule as recited in claim 17, further comprising a high pressureturbine, with each of a low pressure turbine and the high pressureturbine driving a compressor rotor.
 23. The fan drive gear module asrecited in claim 22, wherein the speed reduction device is positionedintermediate a compressor rotor driven by the low pressure turbine and afan section.
 24. The fan drive gear module as recited in claim 22,wherein the speed reduction device is positioned intermediate the lowpressure turbine and the compressor rotor driven by the low pressureturbine.